• Proceedings of the Korean Society of Precision Engineering Conference
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• 2004.05a
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• pp.66-66
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• 2004

중량은 항공기의 성능을 좌우하는 주요 변수로써, 최소의 중량 달성이 100여 년 전 라이트 형제가 최초의 비행을 성공한 후, 항공기 개발자들의 주요 관심사였다. 일반 구조물의 최적화를 통한 설계는 구조물의 강도를 유지하면서 중량절감, 비용절감을 위해 널리 사용되고 있다. 하지만, 현재 항공기 구조물에 대해서는 아직 최적화론 통한 설계가 널리 적용되지 않고 있다. 그래서, 항공기에서는 중량절감을 위해 가벼우면서 강도를 큰 알루미늄이 주재료로 사용되고 있다.(중략)

• Journal of Aerospace System Engineering
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• v.15 no.5
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• pp.106-112
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• 2021

In this study, we performed a design and structural analysis of a blade-shaped antenna installation on the rear fuselage of a rotary wing aircraft operated by the military. When the structure is damaged while the aircraft is in operation, it is separated from the aircraft main structure and may collide with the rotor or blades to cause the aircraft to crash. Therefore, structural safety for the modified structure must be secured. The design requirement for the newly installed modified part were established, and the load condition was constructed by applying the load that may occur in the aircraft after the modification. Structure safety for the modified structure was secured by performing structure analysis. To analysis stress and deformation of aircraft structure, we developed finite element model and verified it by using hand calculation method. We confirmed the safety of the modified structure through the final structural integrity analysis.

• Composites Research
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• v.25 no.4
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• pp.105-109
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• 2012

Recently, development of a small aircraft has been carried out for the BASA(Bilateral Aviation Safety Agreement) program in Korea. This aircraft adopted all composite structures for environmental friendly by low fuel consumption due to its lightness behavior. However the composite structure has disadvantage which is very weak against impact damages. Therefore, damage allowable design of aircraft structure must be performed considering compressive fracture strength. This point is very important for certification of composite structure aircraft. In this paper, it is performed the research on damage tolerance of thick laminate adopting aircraft structure. The damage tolerance of three different types of thick laminates such as no damage, open hole and impact damage is evaluated under compression loading.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.21 no.7
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• pp.638-644
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• 2020

This study examines cracks that occur under the load of an aircraft. The life of aircraft vulnerability structures was analyzed and structural fitting improvements were made. Structural integrity and safety have been achieved through preemptive life expectancy and life management of aircraft structures. The crack size inspection capability of the aircraft under analysis is 0.03inch, compared with 0.032inch, which is the lowest of the three vulnerable parts. In addition, the fatigue life analysis results in approximately 1450 operating hours, the lowest of the three vulnerable parts relative to the aircraft's required life of more than 15000 operating hours, which increased the repeat count of the aircraft's initial and re-inspection times, and hence raised the resulting costs and manpower consumption. Finally, the features were improved through structural fitting of the identified three weak parts. The lowest critical crack size was secured at 0.13 through increased structural resistance to generated cracks and increased aircraft safety. The lowest structural fatigue life for cracks occurring during aircraft operation is 25000 operating hours, which are analyzed above the required structural life, resulting in more optimized improvements than the repair costs and excessive fitting range caused by cracks and fractures.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.21 no.6
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• pp.448-454
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• 2020

This study aims to determine the cause of structural defects occurring during aircraft operations and to verify the structural integrity of the improved features. The fracture plane was analyzed to verify the characteristics of the cracks and the fatigue failure leading to the final fracture was determined by the progress of the cracks by the repeated load. During aircraft operations, the comparative analysis of the load measurement data at the cracks with the aircraft design load determined that the measured load was not at the level of 30% of the design to be capable of being damaged. A gap analysis resulted in a significant stress of approximately 32 ksi at the crack site. Pre-Load testing also confirmed that the M.S. was reduced by more than 50% from +0.71 to +0.43, resulting in a sharp increase in aircraft load and the possibility of cracking when combined. Thus, structural reinforcement and the removal of the gap for aircraft cracking sites improved the defect. Based on the structural strength analysis of the improvement features, the bulkhead has a margin of about +0.88 and the fitting feature is about +0.48 versus allowable stress. In addition, the life analysis results revealed an improvement of approximately 84000 hours.

• Aerospace Industry
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• v.63
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• pp.22-29
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• 1998

지난호에서 미국과 유럽의 항공기 산업의 구조 조정에 대하여 간략한 내용을 살펴본 바 있다. 그중 유럽부문은 민간항공인 에어버스사를 중심으로 한 시장동향과 미국 보잉과의 경쟁 전망을 엮어 보았다. 이번에는 유럽에 있어 각국별 항공기 산업의 재편성에 대하여 국별 개황을 살펴 보기로 한다.

• Aerospace Engineering and Technology
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• v.10 no.2
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• pp.146-153
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• 2011

Some methods of calculation of test load value from design load data were investigated which will be applied at strap installed full-scale airframe of composite aircraft. These methods were applied to left wing of KC-100 composite aircraft and the calculated test load values were compared with each others. Generally since test load values are differently calculated according to each aircraft type and position of straps, all calculation methods mentioned at this study need to be applied and compared to each aircraft. Finally the most appropriate method needs to be selected.

• Proceedings of the Korean Institute of Industrial Safety Conference
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• 2003.10a
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• pp.101-106
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• 2003

항공기 외부 장착물인 POD는 전자전 방해장치(ECM, electronic counter measures)로서 전자방해 장비 및 부분 부품인 전ㆍ후방 러그(lug)와 외부 케이스로 이루어져 있다. POD는 항공기 외부 동체 하단부 및 파일런(pylon)에 장착되어 작동하므로 항공기의 운용중 이륙부터 착륙간의 기동에 의한 피로하중을 주로 받게 되므로 POD 부품들에 대한 구조 건전성 확보하기 위해서는 MIL-STD-1530의 요구에 따라 내구성(durability) 및 손상허용설계(damage tolerance design) 의 평가가 요구되고 있다.(중략)

• Proceedings of the Korean Institute of Industrial Safety Conference
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• 2001.11a
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• pp.118-123
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• 2001

항공기 운항 중에 손상된 부품의 교환이나 수리는 항공기 운항 시 안전 유지와 고가의 항공기 수명연장을 위하여 필수적이다. 여러 가지 손상이나 고장 중에서 균열은 기계적 체결요소(리벳, 볼트 등)나 패칭에 의하여 효과적으로 보수될 수 있다. 항공기 동체 보수 시 접착제 접합을 사용한 복합재료 패칭은 하부구조에 손상이 없이 균열진전을 줄일 수 있고 구멍의 생성으로 인한 응력 집중을 제거할 수 있으며 접합된 면을 외부로부터 차단시켜 부식을 방지하는 효과가 있어 많이 사용된다 특히 고강도 복합재료와 접착제의 개발로 인하여 손상된 구조의 보수를 위한 복합재료 패칭의 사용은 더욱 증가되고 있다.(중략)

• Journal of Aerospace System Engineering
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• v.12 no.4
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• pp.106-111
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• 2018

Hydrodynamic ram phenomenon could be generated by external threats such as impact and blast in the aircraft. High strain rate deformation caused by the hydrodynamic ram phenomenon is one of the main factors to influence structural survivability. Mechanical properties of composite structure change rapidly under conditions of high strain rate. Therefore, it is necessary to experimentally investigate the influence of strain rates for aircraft structural survivability. In this paper, tensile tests of composite material were conducted for low and high strain rates to investigate the influence of the various strain rates. Tensile modulus increases more compared to tensile strength at high strain rate under hydrodynamic ram condition. Regression analysis was conducted to predict tensile modulus at various strain rates because it is one of the main damaging factors for composite structures under high strain rate conditions. Also, the mechanical properties of composite materials were acquired and analyzed under high strain rate conditions. It is hypothesized that the results from this study would be used for designing aircraft composite structures and evaluation considering structural survivability.